A generalized gas turbine engine (GTE) includes an intake section, a compressor section, a combustion section, a turbine section, and an exhaust section disposed in axial flow series. The compressor section includes one or more compressor stages, and the turbine section includes one or more air turbine stages each joined to a different compressor stage via a rotatable shaft or spool. During operation, the compressor stages rotate to compress air received from the intake section of the GTE. A first portion of the compressed air is directed into an annular combustor mounted within the combustion section, and a second portion of the air is directed through cooling flow passages that flow over and around the combustor. Within the combustion chamber, the compressed air is mixed with fuel and ignited. The air heats rapidly and exits each combustor chamber via an outlet provided through the combustor's downstream end. The air is received by at least one turbine nozzle, which is sealingly coupled to the combustor's downstream end. The turbine nozzle directs the air through the air turbines to drive the rotation of the air turbines, as well as the rotation of the spools and compressor stages coupled thereto. Finally, the air is expelled from the GTE's exhaust section. The power output of the GTE may be utilized in a variety of different manners, depending upon whether the GTE assumes the form of a turbofan, turboprop, turboshaft, or turbojet engine.
The sealing interface between the turbine nozzle and the combustor preferably maximizes the operational lifespan of the GTE while simultaneously minimizing leakage between the turbine nozzle and the combustor. It has, however, proven difficult to design a durable, low leakage combustor-turbine seal interface largely due to the extreme thermal gradients that result from temperature fluctuations in the air exhausted from the combustor, as well as the temperature differentials between the air exhausted from the combustor and the cooler air bypassing the conductor. Such thermal gradients cause thermal distortion and relative movement between the various components of the combustor-turbine seal interface; e.g., between the liner walls and the turbine nozzle, which become relatively hot during combustion, and the engine casing, which remains relatively cool during combustion and which may be fabricated from a low thermal growth material, such as a titanium-based alloy. As a result of thermal distortion, leakage paths may form between mating components even if such components fit closely in a non-distorted, pre-combustion state. Compression seals (e.g., metallic W-seals) may be employed to minimize the formation of such leakage paths; however, such compression seals may also be heated to undesirably high temperatures by the hot air exhausted from the combustor, and the sealing characteristics and strength of the compliant seals can be compromised. Furthermore, if the components of the combustor-turbine seal interface are unable to adequately accommodate such thermal distortion, the combustor-turbine seal interface may experience relatively rapid thermomechanical fatigue and decreases in performance. The GTE may consequently require premature removal from service and repair, resulting in economic loss due to the non-availability of the GTE, as well as direct maintenance costs.
There thus exists an ongoing need to provide a combustor-turbine seal interface that significantly reduces or eliminates leakage between a combustor and a turbine nozzle (or nozzles). Ideally, embodiments of such a combustor-turbine seal interface would include one or more compliant structures that accommodate relative movement between the combustor, the turbine nozzle, and the engine casing to reduce thermomechanical fatigue and increase operational lifespan of combustor-turbine seal interface. It would also be desirable for embodiments of such a combustor-turbine seal interface to promote efficient cooling of the combustor and, perhaps, of the leading edge portion of the turbine nozzle. Lastly, it would be desirable for embodiments of the combustor-turbine seal interface to provide aerodynamically efficient flow paths for the heated air exhausted from the combustor, as well as for the cooler air bypassing the combustor. Other desirable features and characteristics of the present invention will become apparent from the subsequent Detailed Description and the appended Claims, taken in conjunction with the accompanying Drawings and this Background.